This invention relates to gas turbine engines and, more particularly, to the fabrication of the turbine disks and seals, and their protection against oxidation and corrosion.
In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot combustion gases are passed through a turbine mounted on the same shaft. The flow of combustion gas turns the turbine, which turns the shaft and provides power to the compressor and to the fan. In a more complex version of the gas-turbine engine, the compressor and a high-pressure turbine are mounted on one shaft having a first set of turbines, and the fan and a low-pressure turbine are mounted on a separate shaft having a second set of turbines. The hot exhaust gases and the air propelled by the fan flow from the back of the engine, driving it and the aircraft forward. The hotter the combustion and exhaust gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the combustion-gas temperature.
The turbine (sometimes termed a xe2x80x9cturbine rotorxe2x80x9d) includes one or more turbine disks, a number of turbine blades mounted to the turbine disks and extending radially outwardly therefrom into the combustion-gas flow path, and rotating seals that prevent the hot combustion gases from contacting the turbine shaft and related components. The maximum operating temperature of the combustion gas is limited by the materials used in the turbine. Great efforts have been made to increase the temperature capabilities of the turbine blades, resulting in increasing combustion as operating temperatures and increased engine efficiency.
As the maximum operating temperature of the combustion gas increases, the turbine disk and seals are subjected to higher temperatures in the combustion-gas environment. As a result, oxidation and corrosion of the turbine disk and seals have become of greater concern. Alkaline sulfate deposits resulting from the ingested dirt and the sulfur in the combustion gas are a major source of the corrosion, but other elements in the aggressive combustion-and bleed gas environment may also accelerate the corrosion. The oxidation and corrosion damage may lead to premature removal and replacement of the turbine disk and seals unless the damage is reduced or repaired.
The turbine disks and seals for use at the highest operating temperatures are made of nickel-base superalloys selected for good toughness and fatigue resistance. These superalloys are selected for their mechanical properties. They have some resistance to oxidation and corrosion damage, but that resistance is not sufficient to protect them at the operating temperatures that are now being reached.
The current state of the art is to operate the turbine disks and seals without any coatings to protect them against oxidation and corrosion. At the same time, a number of oxidation-resistant and corrosion-resistance coatings have been considered for use on the turbine blades. These available turbine-blade coatings are generally too thick and heavy for use on the turbine disks and seals and also nay adversely affect the fatigue life of the turbine disks and seals. There remains a need for an approach for protecting turbine disks and seals against oxidation and corrosion as the operating-temperature requirements of the turbine disks and seals increase. This need extends to other components of the gas turbine engine as well. The present invention fulfills this need, and further provides related advantages.
The present approach provides an approach for fabricating a nickel-base superalloy component of a gas turbine engine, such as a turbine disk or a seal, and components made thereby. The gas turbine component has improved oxidation and corrosion resistance as compared with conventional gas turbine components. There is very little increased weight and added dimension to the turbine component as a result of utilizing the present approach. The present fabrication approach is economically applied and is environmentally friendly. It is not limited by line-of-sight application procedures, so that otherwise-inaccessible portions of the component may be treated. The protection extends over the entire processed surface area of the component, so that protection is provided even in areas where there may be cracks or discontinuities in other applied coatings.
A method for fabricating a gas turbine component comprises the steps of furnishing a substrate shaped as a gas turbine component, such as a gas turbine disk or a seal, and made of a nickel-base superalloy, and oxidizing the substrate to produce an oxidized substrate having thereon a layer comprising an oxide and having a thickness of at least about 500 Angstroms. The step of oxidizing is performed in an atmosphere that does not contain combustion gas. The oxidized substrate is thereafter placed into service.
This approach may be used in conjunction with a number of additional processing steps. The step of furnishing the substrate may include a step of pre-processing the substrate by machining, peening, and grit blasting. A protective coating may be deposited on the substrate, so that the step of oxidizing produces an oxidized coating. The protective coating may include an element such as aluminum, chromium, silicon, phosphorus, or mixtures thereof.
The oxidizing step may be performed in an air atmosphere, so that there is some formation of nitrides as well. The oxidizing step may be performed in an oxygen-only atmosphere, such as from about 0.2 to about 1000 parts per million of oxygen. In a typical case, the step of oxidizing the substrate includes heating the substrate to a temperature of from about 1200xc2x0 F. to about 1550xc2x0 F., for a time of at least about 2 hours.
A top coating may optionally be deposited on the oxidized substrate after the oxidation but before the oxidized substrate is placed into service.
The present invention involves in-situ formation of an oxide layer, not deposition of a coating or a layer from a separate source. The approach does not involve line-of-sight deposition, so that the entire component is protected without regard to position relative to a source. The oxidation is performed after all forging and other mechanical surface processing of the component to its final shape and surface condition are completed, although subsequent coating that does not disrupt the oxide is permitted. Further mechanical operations after oxidation would disrupt the oxide and render it ineffective.
The oxide layer typically has a thickness of from about 1000 Angstroms to about 6000 Angstroms, so that it adds very little weight or dimension to the component. This thin oxide layer improves the oxidation and corrosion resistance of the component by at least 50 percent as compared with an unprotected component, without adversely affecting the mechanical properties such as strength, toughness, and fatigue resistance. The oxide layer includes oxides of the components of the superalloy, such as chromium, titanium, nickel, cobalt, aluminum, and tantalum, and may also include titanium and other nitrides if the oxidation is performed in air.
An important feature of the present processing is that the oxidation treatment is performed prior to the component entering service, and without combustion gas or other gas containing corrosive agents present. Prior turbine components are oxidized when they enter service and are heated to their operating temperatures, but that oxidation is performed in an environment that includes the combustion products which inhibit the formation of a protective oxide and include compounds such as the sulfides and carbides that contribute to corrosion damage. In that prior approach, the corrosive agents are incorporated into the surface of the turbine component before the oxide has a chance to form in the manner of the present approach.
Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment taken in conjunction with the accompanying drawings, which illustrate, by way of example, the principles of the invention. The scope of the invention is not, however, limited to this preferred embodiment.